EP2954171A1 - Turbine à gaz comprenant un thermoplastique pour lisser les surfaces aérodynamiques - Google Patents

Turbine à gaz comprenant un thermoplastique pour lisser les surfaces aérodynamiques

Info

Publication number
EP2954171A1
EP2954171A1 EP14749516.2A EP14749516A EP2954171A1 EP 2954171 A1 EP2954171 A1 EP 2954171A1 EP 14749516 A EP14749516 A EP 14749516A EP 2954171 A1 EP2954171 A1 EP 2954171A1
Authority
EP
European Patent Office
Prior art keywords
gap
set forth
gas turbine
turbine engine
vane
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP14749516.2A
Other languages
German (de)
English (en)
Other versions
EP2954171A4 (fr
Inventor
Thomas B. HYATT
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2954171A1 publication Critical patent/EP2954171A1/fr
Publication of EP2954171A4 publication Critical patent/EP2954171A4/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B05—SPRAYING OR ATOMISING IN GENERAL; APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D—PROCESSES FOR APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D1/00—Processes for applying liquids or other fluent materials
    • B05D1/26—Processes for applying liquids or other fluent materials performed by applying the liquid or other fluent material from an outlet device in contact with, or almost in contact with, the surface
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B05—SPRAYING OR ATOMISING IN GENERAL; APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D—PROCESSES FOR APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D3/00—Pretreatment of surfaces to which liquids or other fluent materials are to be applied; After-treatment of applied coatings, e.g. intermediate treating of an applied coating preparatory to subsequent applications of liquids or other fluent materials
    • B05D3/12—Pretreatment of surfaces to which liquids or other fluent materials are to be applied; After-treatment of applied coatings, e.g. intermediate treating of an applied coating preparatory to subsequent applications of liquids or other fluent materials by mechanical means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
    • F05D2300/433—Polyamides, e.g. NYLON
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
    • F05D2300/436—Polyetherketones, e.g. PEEK

Definitions

  • This application relates to a method and apparatus wherein thermoplastic is deposited into areas of a gas flow path for a gas turbine engine to provide a smooth aerodynamic surface.
  • Gas turbine engines typically include a fan delivering air into a bypass duct, and into a core engine.
  • a compressor sits in the core engine and receives the air flow. Compressed air is passed into a combustor where it is mixed with fuel and ignited, and products of this combustion pass downstream over turbine rotors driving them to rotate.
  • All of the surfaces within the gas turbine engine desirably have aerodynamic efficient shapes.
  • vanes are mounted to guide the air downstream of the fan.
  • the vanes tend to be bolted into an outer housing, and spaced from other housings. In such structures, there are gaps. The gaps can reduce the efficiency of the overall engine, and thus is desirable to smooth these surfaces.
  • a gas turbine engine has a surface configured for being in a gas flow path, the surface having at least one structural member defining a gap.
  • a thermoplastic is deposited into the gap to smooth the surface, whereby the surface is aerodynamically and mechanically smoothly continuous over a gap area.
  • the surface has at least two structural members spaced in an area defining the gap.
  • the gap is between a platform of a vane, and a spaced housing.
  • a second gap surrounds the head of a securement member.
  • a vane extends between a pair of inner and outer wall surfaces, and has platforms attached to each of the inner and outer wall surfaces.
  • the gap includes recesses around a head of a securement member which secures the inner and outer platforms to associated housings.
  • the gap also includes a space between both the inner and outer platforms and an associated housing.
  • the vane sits in a bypass duct.
  • a vane extends between a pair of inner and outer wall surfaces, and has platforms attached to each of the inner and outer wall surfaces.
  • the gap includes recesses around a head of a securement member which secures the inner and outer platforms to associated housings.
  • a method of smoothing an aerodynamic surface in a gas turbine engine includes depositing a thermoplastic into a gap in a surface configured for being in a gas flow path, the surface including at least one structural member defining a gap.
  • the surface is smoothed to remove excess thermoplastic to provide better aerodynamic efficiency whereby the surface aerodynamically and smoothly continuous over a gap area.
  • the surface includes at least two structural members spaced in the gap area.
  • the gap is between a platform of a vane, and a spaced housing.
  • the gap surrounds the head of a securement member.
  • a vane extends between a pair of inner and outer wall surfaces, and has platforms attached to each of the inner and outer wall surfaces. The gap includes recesses around a head of the securement member which secures the inner and outer platforms to associated housings.
  • the gap also includes a space between both the inner and outer platforms and an associated housing.
  • the vane sits in a bypass duct.
  • the gap surrounds the head of a securement member.
  • F3 ⁇ 4 *ure 1 schematically shows a gas turbine engine.
  • F3 ⁇ 4 *ure 2 shows a vane mounted in a bypass duct.
  • F3 ⁇ 4 *ure 3A shows a first problematic location.
  • F3 ⁇ 4 *ure 3B shows the invention applied to the first problem area.
  • F3 ⁇ 4 *ure 4A shows a second problem area in a gas turbine engine.
  • Figure 4B shows the invention applied to the second problem area.
  • F3 ⁇ 4 *ure 5A shows a first step in depositing thermoplastic.
  • F3 ⁇ 4 *ure 5B shows a final step.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B in a bypass duct defined within
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the engine 20 bypass ratio is greater than about ten (10: 1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5: 1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5: 1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet.
  • the flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
  • "Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] 0'5 .
  • the "Low corrected fan tip speed” as disclosed herein according to one non- limiting embodiment is less than about 1150 ft / second.
  • Figure 2 shows a fan rotor 98 delivering bypass air E downstream into a bypass duct where it encounters a vane 100.
  • This may be part of an engine such as shown in Figure 1.
  • the vane 100 is mounted at an inner platform 102 and at an outer platform 104.
  • the outer platform 104 is spaced by a space 110 from an associated housing member 151.
  • a bolt 113 secures the platform 104 to another housing 150.
  • Figure 4A shows a second problematic area wherein the inner platform 102 is spaced by a gap 108 from a forward housing member 106.
  • a bolt 109 secures the platform 102 to a second housing member 103.
  • bolts 113 and 109 are shown, other securement members may be used.
  • Figure 3B shows a material 210 that has been deposited into the gap 110, and material 211 filling the gap 111.
  • Figure 4B shows materials at 208 and 207 filling the prior gaps 108 and 107.
  • Figure 5A shows the material which is deposited to fill the gaps 210, 211, 207 and 208, a thermoplastic.
  • Thermoplastics are known that have melting temperatures well above the operating temperatures that would be seen in the bypass duct, as an example. There are commercially available systems which can deposit the thermoplastic into the gap.
  • a simple tool such as a hot glue gun 300, can melt the thermoplastic such that it flows as shown in 301 into the recess about the head 111 of the bolt 109, as an example.
  • the same process can be utilized at the other areas.
  • Figure 5B shows a subsequent step, wherein a warm putty knife or other tool 310 is utilized to smooth off the surface such that the final smooth shape such as shown in Figure 4B is reached.
  • a method of smoothing an aerodynamic surface in a gas turbine engine 20 includes the steps of depositing a thermoplastic into a gap 210, 211, 207, or 208 in a surface that will be part of a gas flow path when the gas turbine engine is operated.
  • the surface has at least two structural members spaced by the gap.
  • the surface is smoothed 310 to remove excess thermoplastic to provide better aerodynamic efficiency.
  • a gas turbine engine 20 has a surface configured for being in a gas flow path.
  • the surface has at least two structural members spaced in an area defined by a gap 210, 211, 207 or 208.
  • a thermoplastic is deposited into the gap to smooth the surface, whereby the surface is aerodynamically and mechanically smoothly continuous over the gap area.
  • the "structural members” could be the platform 104 and housing member 151, the platform 102 and housing member 106, or the bolts 109/113 and their associated platform.
  • the term “structural members” can extend to many other components that may be found within a gas turbine engine.
  • the term “structural” should not be interpreted to imply load bearing, but rather should be interpreted broadly.
  • the disclosed embodiments show a gap formed between two structural members, this application may extend to a gap formed within a single structural member.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne une turbine à gaz comprenant une surface conçue pour un trajet d'écoulement de gaz. La surface comprend au moins un élément structurel définissant un jour. Un thermoplastique est déposé dans le jour afin de lisser la surface, la surface étant ainsi aérodynamiquement et mécaniquement lisse et continue sur la région du jour. L'invention concerne également un procédé.
EP14749516.2A 2013-02-10 2014-02-04 Turbine à gaz comprenant un thermoplastique pour lisser les surfaces aérodynamiques Withdrawn EP2954171A4 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201361762909P 2013-02-10 2013-02-10
PCT/US2014/014541 WO2014123838A1 (fr) 2013-02-10 2014-02-04 Turbine à gaz comprenant un thermoplastique pour lisser les surfaces aérodynamiques

Publications (2)

Publication Number Publication Date
EP2954171A1 true EP2954171A1 (fr) 2015-12-16
EP2954171A4 EP2954171A4 (fr) 2016-07-06

Family

ID=51300067

Family Applications (1)

Application Number Title Priority Date Filing Date
EP14749516.2A Withdrawn EP2954171A4 (fr) 2013-02-10 2014-02-04 Turbine à gaz comprenant un thermoplastique pour lisser les surfaces aérodynamiques

Country Status (3)

Country Link
US (1) US20150369066A1 (fr)
EP (1) EP2954171A4 (fr)
WO (1) WO2014123838A1 (fr)

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5074752A (en) * 1990-08-06 1991-12-24 General Electric Company Gas turbine outlet guide vane mounting assembly
US5494404A (en) * 1993-12-22 1996-02-27 Alliedsignal Inc. Insertable stator vane assembly
US6619917B2 (en) * 2000-12-19 2003-09-16 United Technologies Corporation Machined fan exit guide vane attachment pockets for use in a gas turbine
GB2400415B (en) * 2003-04-11 2006-03-08 Rolls Royce Plc Vane mounting
US7575702B2 (en) * 2004-04-29 2009-08-18 The Boeing Company Pinmat gap filler
GB2427900B (en) * 2005-07-02 2007-10-10 Rolls Royce Plc Vane support in a gas turbine engine
US7721996B2 (en) * 2007-11-13 2010-05-25 The Boeing Company Fabrication and installation of preformed dielectric inserts for lightning strike protection
GB0905729D0 (en) * 2009-04-03 2009-05-20 Rolls Royce Plc Stator vane assembly
US8322991B2 (en) * 2009-04-10 2012-12-04 Rolls-Royce Corporation Balance weight

Also Published As

Publication number Publication date
US20150369066A1 (en) 2015-12-24
WO2014123838A1 (fr) 2014-08-14
EP2954171A4 (fr) 2016-07-06

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